Method for making a pin reinforced, crack resistant fiber reinforced composite article

ABSTRACT

A composite article, for example a blading member of a gas turbine engine, comprising a plurality of stacked layers of in-plane reinforcing fibers bonded together with a matrix resin is provided with enhanced resistance to impact cracking, material loss and/or delamination through use of a plurality of spaced apart reinforcing pins disposed into the article at an angle to the stacked layers, in one form disposed in a selected article region generally to resist strain energy generated during operation of the region. In another form, enhanced resistance is provided through the combination of a matrix resin including properties comprising a tensile strain property of at least 5% and a K 1c  toughness of at least about 850 psi-inch 1/2 , and a plurality of spaced apart reinforcing pins disposed into the article at an angle to the stacked layers. A method for making such a composite article with such resin comprises stacking the layers of in-plane reinforcing fibers into a shape and then inserting the reinforcing pins shape. The shape is cured with a matrix resin about the in-plane fibers of the stacked layers and about the reinforcing pins.

BACKGROUND OF THE INVENTION

This invention relates to a fiber reinforced composite article andmethod for making such article. More particularly it relates to such anarticle and method, for example a composite blading member including anairfoil, having generally angled or transverse pin-type reinforcingmembers in a toughened and enhanced resin matrix.

Components for sections of gas turbine engines, for example a fan and/ora compressor, operating at relatively lower temperatures than sectionsdownstream of the combustion section have been made of resin matrixcomposites including stacked, laminated layers. Generally such primarilynon-metallic composite structures, which replaced heavier predominantlymetal structures, include superimposed layers, sometimes called plies,reinforced with fibers substantially in the plane of the layer. As usedherein, fibers include within it meaning filaments in a variety ofconfigurations and lay-up directions, sometimes about a core and/or withlocal metal reinforcement or surface shielding. For elevated temperatureapplications, a variety of materials are used for such fibers, includingcarbon, graphite, glass, metals (forms of which sometimes are calledboron fibers), etc., as is well known in the art. Typical examples ofsuch components made primarily of non-metallic composites are reportedin such U.S. patents as No. 3,892,612—Carlson et al. (patented Jul. 1,1975); No. 4,022,547—Stanley (patented May 10, 1977); No.5,279,892—Baldwin et al. (patented Jan. 18, 1994); No. 5,308,228—Benoitet al. (patented May 3, 1994); and No. 5,375,978—Evans et al. (patentedDec. 27, 1994).

As has been discussed in detail in such patents as the above-identifiedEvans et al. patent, such non-metallic composites in an aircraft gasturbine engine are subject to damage from ingestion into the engine andimpact on components of foreign objects. Such objects can be airborne ordrawn into the engine inlet. These include various types and sizes ofbirds as well as inanimate objects such as hailstones, sand, land ice,and runway debris. Impact damage to the airfoil of blading members,including fan and compressor blades, as well as damage to strut typemembers in the air stream, has been observed to cause loss of materialand/or delamination of the stacked layers. Such a condition in arotating blade can cause the engine to become unbalanced resulting inpotentially severe, detrimental vibration.

The above identified and other prior art have reported variousarrangements and structures to avoid such material loss and/ordelamination of layers. Some arrangements, for example U.S. Pat. No.3,834,832—Mallinder et al. (patented Sep. 10, 1974) and theabove-identified Benoit et al. patent, include use of seams or fasteningdevices disposed transversely through a reinforced resin matrix. Theirpurpose is to avoid delamination of laminated composite structuresusing, as the composite matrix, ordinary commercial resin systems havingthe typical relatively low toughness and tensile strain properties. Ithas been observed, however, that disposition of such ordinary transversereinforcement with such ordinary resin systems in modern gas turbineengine blading members such as the airfoil of a fan blade and/or withoutregard to what commonly is referred to in the art as strain energydeveloped in different portions of a blade airfoil, can result in theabove-described type of damage, including delamination and/or materialloss. Such damage can reduce the operating integrity and life of acomposite article.

BRIEF SUMMARY OF THE INVENTION

The present invention, in one form, provides a composite articlecomprising a plurality of stacked layers of reinforcing fibers,preferably disposed in general alignment substantially unidirectionallywith one another in a layer. Such aligned reinforcing fibers, sometimescalled in-plane fibers, generally define at least a portion of the planeof a layer. The layers are bonded together with a matrix resin. Aplurality of spaced apart reinforcing pins are disposed into the articlein at least one selected region, in one embodiment across a selectedregion at a density that generally balances or resists a particularamount of strain energy developed during operation in the selectedregion. In another embodiment, the article that includes the stackedlayers of such arrangement of reinforcing in-plane fibers further istoughened and provided with enhanced resistance to cracking and layerdelamination. That is accomplished through use of a matrix resin thatincludes properties comprising a tensile strain property of greater than5% and a K_(1c) toughness of at least about 850 psi-inch^(1/2) incombination with a plurality of particular additional reinforcingmembers. Such additional reinforcing members, herein referred to forconvenience as pins, are disposed into the article in at least oneselected article region at an angle, for example generally transverse,to the planes of the stacked layers at a selected density, preferablysubstantially uniformly, within the selected region. In one form, thepin comprises a bundle of a plurality of filaments impregnated with aresin as a matrix, preferably substantially completely through thestacked layers.

In another form, the present invention provides a method for making sucha composite article comprising providing the plurality of layers ofaligned reinforcing fibers, in one embodiment impregnated with theabove-defined matrix resin in a partially cured condition. Sometimessuch a partially cured member is referred to as being in the “greenstate” or as a “prepreg”. According to a form of the method of thepresent invention, the layers of fibers are stacked one upon anotherinto a stack of layers. Then the pins are inserted appropriately intothe stack of layers. Thereafter, the matrix resin is cured about thelayers and the pins.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagrammatic side view of a turbine engine blading member,such as a fan blade, including a fiber reinforced composite airfoil.

FIG. 2 is an enlarged, diagrammatic, fragmentary, partially sectionalview of a portion through a thickness of the airfoil of the compositeairfoil of FIG. 1 showing stacked, fiber reinforced layers andreinforcing pins disposed generally transversely to the layers.

FIG. 3 is an enlarged, diagrammatic, fragmentary, sectional view of apin comprising a bundle of filaments impregnated with a resin matrix.

FIG. 4 is a diagrammatic side view of a blading member as in FIG. 1showing a plurality of selected airfoil regions each including a densityof reinforcing pins selected to compensate or as a function of thestrain energy developed during operation in a particular region.

DETAILED DESCRIPTION OF THE INVENTION

In general, disposition of in-plane fiber reinforced, stacked layerswithin a cured resin matrix has provided some strength and someresistance to material loss in such articles as gas turbine engineblading members. A typical example is an aircraft gas turbine engine fanblade, known in the art to be of complex shape, twist, thickness, etc.Such composite structure is lighter in weight than a comparable metalarticle. Therefore, use of a reinforced composite article to replace ametal article has contributed to improvement in operation of a gasturbine engine. However, because such a structure includes stackedlayers or laminations; an impact on the article, typically on theairfoil of a blading member, can cause the layers to separate ordelaminate. An impact on such a laminated article including a matrix ofa commonly used resin having relatively low toughness and tensile strainproperties has been observed to result in damage including delaminationand/or material loss not only to the airfoil but also to the blade baseand shank.

Damage from impacts has been observed on an article including suchcommonly used resin matrix systems, even though the article has includedthe type of known transverse additional reinforcement, for example asdescribed above in connection with the Mallinder et al. and the Benoitet al. patents. Also, it has been observed that pinning, stapling, orstitching substantially uniformly across an entire blading memberairfoil, without regard to variations in strain energy developed indifferent regions during operation, can result in more damage to ordelamination of the airfoil than one without any additional angled ortransverse reinforcement. It is believed that such damage in theairfoil, typically a gas turbine engine fan or compressor blade airfoil,using such known pinning can result from different operating stresses orstrain energy generated between airfoil regions during operation. Suchstrain energy amounts, as is well known in the gas turbine engine art,are a function of loading and forces on the airfoil during engineoperation. Angled or transverse pinning only in one region of an airfoilor uniformly across the entire airfoil, without regard to variability ofstrain energy developed, can tend to drive the strain energy toward aregion of the airfoil not adequately reinforced to resist such energy.This can result in damage or failure such as delamination in the regioninto which such excess strain energy has been driven. Therefore, it hasbeen recognized, according to forms of the present invention, that adifferent, selected density or amount of such additional reinforcementis required in and between selected regions of an airfoil, as a functionof the desired or required resistance to strain energy developed duringoperation.

As used herein, the term “strain energy”, as is well known in the gasturbine engine art, means the kinetic or dynamic energy developed in ablade airfoil during operation. The result is an amount of deflectionand straining of the airfoil in a particular region of the airfoil. Forexample, thinner portions of the airfoil, such as at the airfoil tip andthe leading and trailing edges, develop less strain energy duringoperation. According to forms of the present invention, such thinnerregions require lower densities of pinning type of reinforcement than doother regions of the airfoil located inwardly of such thinner regions.In an airfoil, there can be a plurality of regions about the airfoilrequiring different amounts of such additional reinforcement dependingon the strain energy developed in each region during operation.Disposition of such additional reinforcement, for example pinning,according to forms of the present invention, is selected about theairfoil substantially to balance or compensate for the variability ofstrain energy developed in the airfoil during operation.

In a preferred form of the present invention, a plurality of selectedregions together substantially cover the entire airfoil generallybetween, but not necessarily including, the limits of the leading andtrailing edge portions and the tip and base portions. However, thedensity of pinning varies between adjacent or contiguous regions tobalance or resist strain energy developed in a region. For example ashas been stated, a region adjacent or in the vicinity of an airfoil tip,a leading edge, or a training edge generally will require a lowerdensity or amount of pinning than a region or regions generally in themiddle portion of the airfoil. In a preferred form, pinning within aregion is disposed substantially uniformly within that region. Sucharrangement, in some embodiments, can eliminate, further, variations instress energy even within a region. As was mentioned, it is one objectof the present invention to disperse strain energy substantiallyuniformly about an airfoil that includes a plurality of regions thatdevelop different strain energies.

Use of the above-defined toughened, high tensile strain resin as amatrix of a fiber reinforced composite article comprising stacked layersof aligned, in-plane reinforcing fibers, in combination with additionalangled, preferably transverse, reinforcing pin bundles, provides anarticle resistant to cracking, crack propagation, and layerdelamination. As it relates to an airfoil of a gas turbine engineblading member, such a laminated construction with the toughened, hightensile strain resin produces a solid structure with proper orthotropicproperties which meets centrifugal and gas loading requirements. Inaddition, the angled, preferably transverse, pin reinforcement disposedwithin at least one selected region of an article provides improvedout-of-plane strength through the thickness of the article, such as theblading member airfoil, necessary to withstand higher impact loads thatcan be experienced during service operation and to inhibit propagationof delamination resulting from such impacts. Furthermore, forms of thepresent invention as an airfoil of a blading member of a gas turbineengine include a plurality of selected regions together substantiallycovering the airfoil, with the density of additional reinforcement beingdifferent between regions as a function of the strain energy developedin the airfoil.

The present invention will be more fully understood by reference to thedrawings. FIG. 1 is a diagrammatic side view of a typical gas turbineengine composite, laminated, fiber reinforced blading member, in theform of a fan blade, shown generally at 10, for example in a form shownin the prior art. Blade 10 includes an airfoil 12, a base 14, a firstside surface 16, a second surface side 18, and an airfoil tip 20.Sometimes first and second side surfaces 16 and 18 are referred to asthe pressure and suction sides of an airfoil. Airfoil 12 includes athickness 22, shown in more detail in FIG. 2, and which can vary acrossairfoil 12 as a function of its design. Airfoil 12 includes additionalreinforcing pins, some of which are shown at 24, disposed within anentire airfoil surface or region 26 included within broken line 28.Region 26 is spaced apart from airfoil portions immediately at theairfoil tip, edges and base portions of the airfoil. However, as will bediscussed in connection with similar FIG. 4, which represents,diagrammatically, an airfoil region arrangement according to a form ofthe present invention, generally airfoil 12 includes a plurality ofcontiguous regions generally about the airfoil.

In the embodiment of FIG. 1, surface region 26 substantially covers theentire surfaces 16 and 18 of airfoil 12, with reinforcing pins 24disposed into airfoil 12 generally transversely to airfoil surfaceregion 26. In the form of FIGS. 1 and 2, reinforcing pins 24 extendsubstantially completely through airfoil 12, from surface 16 to surface18.

FIG. 2 is an enlarged, diagrammatic, fragmentary, partially sectionalview through a portion of thickness 22 of airfoil 12 within region 26.Reinforcing pins 24, one shown as protruding from the fragmentarysection, are disposed within and, in this example, substantiallytransversely to a plurality of typical stacked, fiber reinforcedcomposite planes or layers 30 in airfoil 12.

FIG. 3 is an enlarged, diagrammatic, fragmentary sectional view of areinforcing pin shown generally at 24. Pin 24 comprises a bundle orplurality of generally aligned filaments 32, for example made of carbon,graphite, glass, metal, or their mixtures, held in a resin matrix 34about filaments 32.

FIG. 4 is a diagrammatic side view of a typical gas turbine enginecomposite, laminated fan blade 10 generally as shown in FIG. 1. However,the embodiment of FIG. 4, as one diagrammatic embodiment of the presentinvention in respect to gas turbine engine blading members, includes aplurality of selected surface regions of the airfoil. A first region 36is generally a “U” shaped region extending about airfoil 12 justinwardly from leading edge 38, tip 20, and trailing edge 40. Firstregion 36 is contiguous with a second region 42, also generally of a “U”shape, inwardly of first region 36. Broken line 44 generally representsa boundary area between first region 36 and second region 32. A thirdregion 46, is shown inwardly of second region 42, contiguous with secondregion 42. Broken line 48 generally represents a boundary area betweensecond region 42 and third region 46. A fourth region 50, in theembodiment of FIG. 4, is disposed inwardly of third region 46, withbroken line 52 representing a boundary area between third region 46 andfourth region 50. It should be understood that the regions of aparticular airfoil, for example of a gas turbine engine fan orcompressor blade airfoil, can vary from the diagrammatic embodiment ofFIG. 4, with the different regions disposed about the airfoil as afunction of the variation in strain energy developed in a particulardesign of airfoil during operation.

As shown in embodiment of FIG. 4, the densities of pins 24 withinregions 36, 42, 46 and 50 generally are different between contiguousregions as a function of the strain energy developed during operation ofthe airfoil, substantially to balance the airfoil strain energygenerated during operation of the airfoil. In that embodiment, thedensity of reinforcing pins in each region increases between regionsfrom first region 36 through fourth region 50. This is a diagrammaticexample of disposition of such pins to balance, or resist transfer of,airfoil strain energy during operation

Density as used herein in respect to the pins in a region means theratio of the sum of the surface areas of all pins in a region to thetotal surface area of the region. It should be understood that the crosssectional area of particular pins can be different from one anotherwhile substantially maintaining the selected pin density within aregion. For example, a larger number of pins each with a relativelysmaller cross sectional area can provide improved surface contact with asurrounding matrix.

Within each such region, it is preferred that pins 24 be disposedsubstantially uniformly within that region further to avoid creation ofstress concentrations within that region. As shown in the embodiment ofFIG. 4, the density of pins 24 in region 50, generally in a middleportion of airfoil 12, is greater than in either regions 36, 42 and 46,with pins 24 substantially covering each region to avoid stressconcentrations within the respective region. As the present inventionrelates to gas turbine engine airfoils, it is preferred that the densityof the pins in a region be in the range of about ½ to about 5%, morespecifically in the range of about ½ to about 2%. It has been observedthat a pin density of less than about ½% in a region is insufficient toprovide required additional reinforcement and resistance to strainenergy. In addition, a pin density of greater than about 5% in a regionprovides excessive resistance to strain energy or stiffness that canlead to what sometimes is referred to in the art as notch sensitivity.More particularly in a gas turbine engine laminated composite fan orcompressor blade, a pin density in a region is preferred to be in therange of about ½-2%.

One series of evaluations associated with the present invention wasconducted to compare ordinary, commercially available heat curable epoxyresin systems, currently used in the art as matrix resins for laminatedcomposite articles, with the heat curable epoxy forms of theabove-defined type of toughened resin, included in forms of the presentinvention. Such a comparison was made with the matrix resin cured instacked layers of aligned reinforcing fibers. Currently used cured resinsystems, one example of which commercially is available as Dow ChemicalTACTIX 123 epoxy resin system, have a tensile strain property of about5% or less, for example in the range of about 2-5%, in combination witha K_(1c) toughness of less than about 850 psi-inch^(1/2), for example inthe range of about 400-500 psi-inch^(1/2). These properties both aretypical of currently used epoxy resin systems.

It was recognized that use of such current resin systems in certainapplications, alone and without additional angled or transversereinforcement, resulted in insufficient resistance to impact damage anddelamination caused by ingestion of such objects as birds, hail, andland ice into the engine. For example, one type of test impacted thecured laminated structure with an object equivalent to a 2.5 pound bird.The lower tensile strain of current systems was insufficient to resistimpact damage or loss of material, and the lower toughness levelprovided too low a threshold at which cracking and layer delaminationcould be initiated. This lower threshold results in unacceptablecomponent matrix loss causing performance and balance conditionsdetrimental to the engine. In contrast according to a form associatedwith the present invention, a matrix, for stacked reinforcing layers, ofa cured epoxy resin system with a tensile strain property of greaterthan 5%, for example about 7% or more, in combination with a toughnessK_(1c) of at least about 850 psi-inch^(1/2), provided resistance todegradation, such as material damage and delamination, from such animpact.

In one specific evaluation series of a form of the present invention, aplurality of shaped layers of substantially aligned, unidirectional,in-plane carbon fiber bundles, commercially available as IM-7 12K towtape from Hexel Company, were pre-impregnated, as is commonly practicedin the art, with the above-defined toughened resin to provide layers 30.In this example, the above-defined high strain, toughened epoxy resinwas one identified as 3M PR520 epoxy resin system. Properties of thatresin system included a tensile strain of about 6.9% and a toughnessK_(1c) of about 1380 psi-inch^(1/2).

Such prepreg layers were disposed in a stack of layers in a supportingfixture, with typical amounts of intermediate wicking felt as used inthe art. While in the fixture, a plurality of reinforcing pins 24 wereinserted into the stack of layers, appropriately in selected densitiesuniformly within selected regions, substantially transversely to theplanes of the layers and generally completely through the stack oflayers, to provide a preform. Each pin comprised a plurality or bundleof carbon filaments 32 held in a matrix 34 of a resin compatible withthe resin matrix of the prepreg layers, in this example an epoxy resinin the fully cured condition. The preform was compacted or debulked, ascommonly is practiced in the art, and disposed in the cavity of a curingmold.

After closing the cavity of the mold, a vacuum was provided in thecavity to remove ambient air from the cavity and from about the preform.The mold and its contents were heated in the range of about 350-400° F.for about 90-120 minutes to cure, concurrently, the resin into a matrixabout the fibers and the resin into a matrix about the bundle offilaments in the reinforcing pins 24. This provided an in-plane fiberand transverse pin reinforced carbon/epoxy composite blade. Aftercuring, the mold and its contents were cooled. When removed from themold cavity, the resulting article was a near net shape molded epoxyresin carbon fiber reinforced composite blade including a molded airfoiland molded base. The cured fiber reinforced composite blade using theabove-defined high tensile strain, toughened epoxy resin as a matrix forthe stacked layers along with the transverse pin reinforcement providedimproved impact capability at the point of impact as well as away fromthe impact site while retaining blade spanwise and chordwise directionalstrength capability.

The result of several bird impact tests of the above-describedconstruction form of the present invention showed a decrease of about54% in the amount of delamination compared with a similar blade withoutpinning according to the present invention. Any delamination of theabove-described blade form of the present invention was confined to theairfoil, whereas delamination of the unpinned blade airfoil delaminatednot only in the airfoil but also in the shank and base portions.

Other series of comparison testing were conducted. In one series, apreferred form of the article of the present invention as gas turbineengine blades, constructed with reinforcement pins protrudingsubstantially completely through the thickness and coveringsubstantially the entire surface of the blade airfoil, was compared withsimilarly constructed articles with no transverse reinforcement. Theresults showed that practice of the present invention greatly improvedthe impact performance over articles without such transversereinforcement. The absence of such transverse reinforcement was noteffective in reducing delamination and hastened the progression ofdelamination and material loss.

Embodiments of the present invention, including a method for making,provide a tough, impact and delamination resistant resin matrixcomposite article including a combination of in-plane fiberreinforcement and additional reinforcement disposed into the article atan angle to the in-plane fibers. One example of such an article is ablading member of a gas turbine engine. Resin properties are tailored toprovide high impact resistance while maintaining, for example in anairfoil, spanwise properties to resist centrifugal and bending loads aswell as chordwise properties to resist gas and torsional bending loads.This combination of enhanced properties and multiply types ofreinforcement improves the overall operating capability of the article.

The present invention has been described in connection with specificexamples and combinations of materials and structures. However, itshould be understood that they are intended to be typical of rather thanin any way limiting on the scope of the invention. Those skilled in thevarious arts involved, for example technology relating to gas turbineengines, to fiber reinforced composite structures, fibers and resins,etc, will understand that the invention is capable of variations andmodifications without departing from the scope of the appended claims.

1-23. (canceled)
 24. A method for making a pin reinforced compositearticle comprising a plurality of stacked layers of in-plane reinforcingfibers bonded together with a matrix resin and defining a thickness ofthe article comprising the steps of: providing a plurality of layers ofin-plane reinforcing fibers impregnated with a partially cured matrixresin that includes properties comprising a tensile capacity of greaterthan 5% and a K_(1c) toughness of at least about 850 psi-inch^(1/2);stacking the layers one upon another into a stack of layers; selectingat least one article region of the stack of layers in which a pluralityof reinforcing pins are to be disposed; and, inserting the plurality ofreinforcing pins into the stack of layers at an angle to the stack oflayers.
 25. The method of claim 24 in which the pins are disposed withina selected region at a density of pins that resists strain energygenerated during operation of the region.
 26. The method of claim 25 inwhich the pins are disposed substantially uniformly across a region. 27.The method of claim 24 in which: the in-plane reinforcing fibers in alayer substantially are aligned with one another; and, the reinforcingpins are inserted substantially completely through the thickness of thearticle.
 28. The method of claim 25 in which: the layers of the in-planereinforcing fibers are stacked in a supporting fixture to provide astack of layers; the reinforcing pins are inserted into the stack oflayers to provide a preform; the preform is placed in a curing moldcavity; the mold cavity is closed; a vacuum is provided within the moldcavity about the preform of the layers and the reinforcing pins; and,the matrix resin is cured about the in-plane fibers and reinforcingpins.
 29. The method of claim 28 in which the partially cured matrixresin is cured at a temperature in the range of about 350-400° F. 30-39.(canceled)